Numerical Simulation For Chemical Reacting Flows With Application To Low Speed Range

It is well known that in a scramjet combustor there exist various complex flows such as shock wave, vortex, ignition and combustion, shock wave/boundary layer interaction and separation flow, etc., where both supersonic flow in the core region and subsonic (low speed) flow within the boundary layer or cavity region are common. On the other hand, an aero-engine combustor has even more complex structure in which the combustion involves strong swirl and recirculation flows and exhibits both compressible and incompressible flow characteristics simultaneously. As numerical simulation has advantages of providing not only flow configuration, but also detailed parameters of the complicated flow fields, it can offer effective way for clarifying complex mechanism as well as the optimization of the engine design. In this work, the complex flow of scramjet and aero-engine combustor were studied after an improvement of numerical simulation. Based on 3-D multi-component Navier-stokes equations, a preconditioning, finite volume method with finite-rate chemical reacting model and high-order AUSM+-up scheme was developed for space discretization and implicit LU-SGS scheme for time marching. These schemes were combined to develop a high efficiency and precision massively parallel CFD numerical software platform for gas fuel chemical reacting flow at full-speed range. The first verification was the simulation of reacting flow with or without preconditioning on sw machine, which confirmed the reliability of the new developed software. Then we conducted a series of numerical simulations of the internal flow of the scramjet and aero-engine combustor and the good agreement with experimental data were found and valuable analyses were performed.The main works can be summarized as follows:1. A massively parallel numerical software platform for simulating from low speed to supersonic flow was developed. Based on the existing 3-D multi-component Navier-stokes equations, together with finite chemical reaction rate model and finite volume method, a preconditioning matrix using variables of pressure, velocity and temperature was applied for the virtual time derivative term of the equations. High-order AUSM+-up scheme was used for space discretization, with implicit LU-SGS scheme for time marching. It was verified that the improved program, which utilized traditional theory of compressible flow, is capable of solving chemical reacting flow at all speeds, which means that the capability of simulating low speed,incompressible flow was greatly enhanced.2. A series of numerical simulations were carried out for a rectangular direct-connected scramjet combustor with different equivalence ratio and fueling distribution, in which the detailed structures of the combustion flow fields were obtained. The results show that the ignition mechanism and reacting region are affected by equivalence ratio remarkably, which results in the combustion performance accordingly. It was found that for ethene injection ahead of a cavity flame holder, there exist three typical stabilized combustion modes, namely jet-wake recirculation-zone stabilized combustion, cavity recirculation-zone stabilized combustion, jet-wake/cavity recirculation combination stabilized combustion. The mixing and combustion characteristics with a planar separator were explored. The results show that there exhibit strong unsteady behaviors in the supersonic mixing combustion flow fields. The unsteady flow in the upper side can cause disturbance to the lower part of the flow field and generate perturbed back-pressure which may disturb the inlet flow.3. The combustion flow of a single-head aero-engine was simulated and the results were verified with experimental data reasonably well, from which we got the confidence that the previous code for simulating high speed compressible flow can be improved to solve low speed combustion flow. The numerical analysis of the single-head aero-engine flow shows an attractive potential that the developed code is useful for the optimal design of aero-engine combustor as well as the better understanding of complex engine flow.In summary, the improved numerical method can be used to investigate the combustion flow of both scramjet and aero-engine combustors with relatively high efficiency and accuracy, which is helpful for offering not only the detailed information of complex flow field, but also in-depth performance analysis.